The present invention relates generally to systems for radiating heat from a spacecraft, orbiting vehicle or the like, and more particularly to an improved high efficiency, deployable two-phase heat exchange system.
In the operation of spacecraft in earth orbit, high peak power generation during certain portions of the duty cycle ordinarily require that a radiator system for rejecting waste heat generated aboard the spacecraft be sized to accommodate large heat loads. A conventional heat radiator system typically comprises a two-phase (liquid/vapor) closed system having large radiating surfaces and takes advantage of a high boiling heat transfer rate of the liquid medium. Such a system is susceptible to damage by invasive laser radiation directed onto the liquid tube or heat radiating surfaces which may cause overpressure and overheating within the system and/or by micrometeroid penetration, which damage may affect onboard operating systems the continued functioning of which requires efficient heat rejection.
Representative two-phase systems include those described by or referenced in U.S. Pat. No. 3,496,995 ("Furlable Heat Exchanger"), U.S. Pat. No. 4,212,347 ("Unfurlable Heat Pipe"), U.S. Pat. No. 4,706,740 ("Ventable Survivable Heat Pipe Vapor Chamber Spacecraft Radiator") and U.S. Pat. No. 4,727,932 ("Expandable Pulse Power Spacecraft Radiator").
Prior art systems such as those just cited are characterized by a single liquid/vapor heat exchange system which may be susceptible to substantial damage by invasive laser irradiation or by micrometeoroid impact, which in turn may threaten the continued operation of systems from which heat is to be removed. Further, most prior art systems are characterized by a high degree of expansion (100 meters or more) in a high heat rejection mode, which presents a substantial cross section for micrometeoroid impact and for detection in potentially hostile situations, and which generates a moment about the vehicle resulting in impaired maneuverability thereof.
Conventional heat exchanger structures applied to a two-subsystem configuration suffer from two fundamental shortcomings. First, those structures that use condensation heat transfer rely on gravity to drain condensate from the heat transfer surface. In near zero gravity, a condensate layer builds up on the heat transfer surface and results in a large temperature drop. Second, conventional structures that would use clamped joints between the two sub-systems are generally massive, and have poor heat transfer efficiency that result in a large temperature drop.
The invention eliminates or reduces in critical importance deficiencies in prior art systems by providing a deployable, two-subsystem based radiator system for radiating heat from an orbiting vehicle comprising a plurality of deployable, retractable panels thermally coupled by variable conductance heat pipes with flexible joints. Each module contains a heat transfer interface at which vapor from the heat transport subsystem condenses directly on the outer walls of the radiator heat pipes having grooved surfaces for facilitating condensation. Vapor and liquid lines of each module are housed inside a micrometeoroid shielded boom configured to be flexible and steerable to promote survivability and to optimize heat exchange efficiency. The two-subsystem structure characteristic of the invention is substantially less vulnerable to hostile or natural threats. Further, separate interfacing heat exchange systems, one for conducting heat away from onboard operating systems and a second for radiating that heat into space, include separate heat exchange media which allows versatility in temperature range of operation of the two subsystems.
It is therefore a principal object of the invention to provide a heat rejection system for an orbiting satellite or like vehicle.
It a further object of the invention to provide a deployable heat radiator having large heat rejection capability.
It is yet a further object of the invention to provide a two-subsystem heat rejection system for an orbiting vehicle including a deployable radiator having minimal cross section for detectability or micrometeoroid impact in the deployed condition.
These and other objects of the invention will become apparent as the detailed description of representative embodiments proceeds.